The present invention relates generally to rocket engines and more particularly to methods of joining a ceramic matrix composite rocket nozzle to a metal manifold.
An increasing amount of composite structures are being utilized in certain aerospace vehicle applications, primarily due to significant weight savings that can be achieved over conventional metallic structures. However, the performance improvements of the composite structures must be carefully balanced with overall cost, and as a result, a combination of composite structures and conventional metallic structures commonly exist within an aerospace vehicle. Accordingly, the composite structures must be joined to the conventional metallic structures for vehicle assembly without significant performance degradations or increased costs.
In aerospace vehicles, a significant portion of the vehicle structure may be subjected to extreme thermal conditions, wherein the structure must be capable of withstanding relatively high aerothermal loads during a variety of conditions. For example, rocket nozzles experience temperatures in excess of 1,000xc2x0 C. (1,832xc2x0 F.) for extended periods of time, and accordingly, the nozzles must comprise materials that can withstand such high temperatures over an extended period. In a known application, the rocket nozzle is a composite structure that comprises carbon fibers within a ceramic matrix in order to withstand the high temperatures. The rocket nozzle must further be joined to a manifold, which is generally a metal material, however, conventional techniques such as brazing are not suitable between the ceramic composite structure and the metal manifold due to high localized temperatures.
The joining of ceramic components to metal components is known in the art and may be accomplished using, for example, pressure insertion, thermal insertion, chill insertion, or a combination of such methods along with soldering or brazing. For example, U.S. Pat. No. 5,365,661 to Mizuno et al. discloses a ceramic-metal joint body that controls circularity of a ceramic member during a joining process. Unfortunately, the ceramic-metal joint body encases the metal member in its entirety, which is not practical in aerospace applications due to constraints on overall vehicle weight and performance.
Yet another method of joining a ceramic component to a metal component is disclosed in U.S. Pat. No. 5,552,670 to Heider et al. Generally, Heider et al. discloses a vacuum-tight seal between a ceramic tube and a metal tube that uses plugs inside the ceramic tube into which the metal tube is sintered. Unfortunately, the sintering occurs at relatively high temperatures that are not suitable for a ceramic composite structures as previously set forth. Further, tensile stresses develop across the joint between the ceramic tube and the metal tube upon cooling from the sintering process due to the larger thermal contraction of metals over that of ceramics. Moreover, the metal tube of Heider et al. is disposed inside the plug, which is disposed inside the ceramic tube, thereby creating a significant surface discontinuity along the outer surfaces thereof. Such surface discontinuities are unacceptable in aerospace applications due to the disruption of aerodynamic flow over the outer surfaces.
Accordingly, there remains a need in the art for a method of joining a ceramic matrix composite to a metal that is cost effective and that can withstand the extreme thermal conditions experienced in, for example, a rocket engine nozzle. A need further exists for a method of joining a ceramic matrix composite rocket nozzle to a metal manifold without performance degradations to the aerospace vehicle.
In one preferred form, the present invention provides a method of joining a ceramic matrix composite conduit to a metal conduit using an insert disposed inside both the ceramic matrix composite conduit and the metal conduit. In one form, the ceramic matrix composite conduit is a rocket nozzle, hereinafter referred to as a ceramic matrix composite rocket nozzle, and the metal conduit is a metal manifold, which have particular application to a rocket engine. It shall be appreciated, however, that the present invention has particular utility for applications other than rocket engines and the reference to rocket engines herein shall not be construed as limiting the scope of the present invention.
The insert is preferably co-processed with the ceramic matrix composite conduit, wherein the insert is placed inside a fiber preform and the fiber preform is infiltrated with a ceramic matrix and subsequently heat treated. Accordingly, the ceramic matrix bonds the insert to the inside of the ceramic matrix composite conduit during the infiltration process.
Furthermore, the insert that joins the ceramic matrix composite conduit to the metal conduit is preferably a silicon nitride or other material that has approximately the same or a smaller coefficient of thermal expansion as the ceramic matrix composite conduit. Generally, the coefficient of thermal expansion is approximately equal to or less than the ceramic matrix composite conduit in order to minimize tensile stresses that would occur between the insert and the ceramic matrix composite conduit during cooling after the infiltration process.
After processing of the ceramic matrix composite conduit, the metal conduit is then placed over the insert and secured thereto, preferably using brazing. Due to the larger coefficient of thermal expansion of the metal conduit over that of the insert, compressive stresses are introduced into the insert while the metal conduit experiences tensile stresses. Such a stress state is tailored to the specific properties of ceramics and metals, wherein ceramic materials exhibit high compressive strength and low tensile strength, while metal materials exhibit high tensile strength. Furhermore, a monolithic ceramic insert is preferred over a composite insert since the compressive strength of ceramic composites are lower than those of monolithic ceramics. As a result, a joint assembly is provided that is cost and weight effective, and which comprises a relatively smooth outer moldline surface for aerospace applications.
Further areas of applicability of the present invention will become apparent from the detailed description provided hereinafter. It should be understood that the detailed description and specific examples, while indicating the preferred embodiment of the invention, are intended for purposes of illustration only and are not intended to limit the scope of the invention.